A method for stability expansion design of a compressor in a gas turbine and an aero-engine

Abstract

The application belongs to the technical field of stability expansion design of a gas turbine and an aero-engine compressor, and particularly relates to a stability expansion design method of a gas turbine and an aero-engine compressor, which comprises the following steps: setting a bleed air pipeline to guide part of the airflow at the outlet of the compressor to the inlet position and re-inject the airflow into the main flow passage of the compressor in the form of a high-speed jet; designing the airflow with a flow rate of more than 10% of the inlet flow rate of the compressor to flow back to the inlet position of the compressor through the bleed air pipeline; and designing a jet slot with a continuously contracted cross section at the inlet position of the compressor to enable the airflow in the bleed air pipeline to be injected into the main flow passage in the form of a high-speed fan-shaped area under the driving of the pressure difference between the inlet and the outlet of the compressor, wherein a small part of the airflow is close to the rotor tip end casing, and most of the airflow is interchanged with the main flow in the rotor blade of the inlet stage and above.

Description

Technical Field

[0001] This application belongs to the technical field of extended stability design of intermediate compressors in gas turbines and aero engines, and specifically relates to a method for extended stability design of intermediate compressors in gas turbines and aero engines. Background Technology

[0002] The compressor is a crucial component in gas turbines and aero engines, and its stability margin directly determines the overall stability of the gas turbine and aero engine.

[0003] To improve the stability margin of compressors in gas turbines and aero engines, current technical solutions include:

[0004] 1) From the perspective of the overall architecture, the single-shaft multi-stage compressor is designed as a dual-rotor or triple-rotor structure, with high-pressure and low-pressure compressors or high-, medium- and low-pressure compressors working together to complete the compression task, in order to improve the uncoordinated operation of the front and rear stages of the single-shaft multi-stage compressor and thus improve the compressor stability margin. However, this will greatly increase the complexity of the overall structure of the gas turbine and aero-engine.

[0005] 2) Optimizing the compressor aerodynamic layout and blades, optimizing the compressor adjustable stator adjustment law, optimizing the compressor blade tip clearance control, and adopting compressor handling casing design can improve the compressor stability margin to a certain extent, but the effect is limited and it is difficult to guarantee the compressor stability margin across the entire speed range.

[0006] 3) Employ interstage or outlet venting technology, such as by installing bypasses between compressor stages or at the outlet. Figure 1 As shown, increasing the equivalent exhaust area significantly increases the flow rate, which can lower the matching point and, under the premise that the instability boundary remains unchanged, significantly expand the stable operating range, thereby improving the compressor stability margin. For transition states with large stability margin requirements, increasing the venting volume can further lower the matching point and expand the stable operating range. However, this technical solution has the following drawbacks:

[0007] a) Directly releasing gas into the environment will generate significant noise, and prolonged release will cause continuous high-frequency noise, reducing the quality of the human-machine environment.

[0008] b) The formula for calculating isentropic compression efficiency is: in, This refers to the isentropic compression efficiency of the compressor. L is the isentropic compression work per unit flow rate of the compressor. K G is the rim work per unit flow rate of the compressor. t G is the compressor inlet gas flow rate. bLet λ be the compressor vent flow rate and λ be the compressor vent ratio. It can be seen that the reduction in compressor efficiency is equal to the vent ratio. When using interstage or outlet vent technology, the vent ratio is large. Generally, the direct vent to the environment accounts for 5% to 10% or even higher of the total flow rate, resulting in a significant loss of compressor efficiency and a significant reduction in the overall thermal cycle efficiency.

[0009] 4) The self-circulating bleed air-jet stabilization technology uses high-pressure gas drawn from the compressor interstage or outlet and jets it at the compressor inlet stage or other critical stage rotor and stator blades considered to limit stability margin. This locally suppresses and delays boundary layer stall separation and other secondary flows, thereby improving the compressor stability margin. In most cases, this technology can ensure that the compressor efficiency does not decrease significantly, and sometimes even increases slightly. However, this technology has the following drawbacks:

[0010] a) The application targets are "single rotor" or "single stage compressor", with occasional two-stage or three-stage compressors, and rarely more-stage compressors. It is significantly different from multi-stage compressors in engineering. The influence of interstage matching factors on the stability boundary is weak or even non-existent. For compressors with complex interstage matching, the engineering guidance is insufficient.

[0011] b) Characterized by small or micro-flow bleed air, with the bleed air ratio typically less than 1% of the compressor intake flow, micro-jet jets improve localized, small-scale flow, such as blade surface stall separation and tip leakage flow, improving the overall flow near the compressor surge point. This allows the iso-speed characteristic line to extend outwards naturally near the surge point, thereby delaying compressor stall and instability. Small or micro-jet jets do not significantly affect the overall compressor flow framework, and the improvement in compressor stability margin is limited, generally with an upper limit of about 10%, which is not easy to achieve.

[0012] This application is made in view of the aforementioned technical deficiencies.

[0013] It should be noted that the above background information is only used to assist in understanding the inventive concept and technical solution of this invention, and it does not necessarily belong to the prior art of this application. In the absence of clear evidence that the above information was disclosed on the filing date of this application, the above background information should not be used to evaluate the novelty and inventiveness of this application. Summary of the Invention

[0014] The purpose of this application is to provide a compressor stability enhancement design method for gas turbines and aero engines to overcome or mitigate at least one of the known technical defects.

[0015] The technical solution of this application is:

[0016] A method for enhancing the stability of compressors in gas turbines and aero engines includes:

[0017] Set up an air intake pipeline to divert part of the airflow from the compressor outlet to the inlet position and re-inject it into the main airflow path of the compressor;

[0018] The design allows more than 10% of the compressor intake airflow to return to the compressor inlet through the bleed air pipeline.

[0019] A continuously contracting circumferential jet channel is designed at the compressor inlet position, so that the airflow in the intake pipe is driven by the pressure difference between the compressor inlet and outlet, and is injected into the main flow channel at high speed in a fan-shaped area. A small part of it is close to the top of the rotor blade casing, while most of it interacts with the main flow in the inlet stage rotor blade and above.

[0020] According to at least one embodiment of this application, the above-described compressor stability enhancement design method for gas turbines and aero engines further includes:

[0021] A throttle valve is installed on the bleed air pipeline to adjust the intensity of the bleed air-jet circulation flow at different speeds. The goal is to meet the instability boundary or margin improvement requirements, with the basic principle of minimizing the circulation intensity and thus reducing compressor losses.

[0022] According to at least one embodiment of this application, the above-described compressor stability enhancement design method for gas turbines and aero engines further includes:

[0023] In a high speed range, including the design speed, where the instability boundary is high enough or the stability margin is large enough, and no further stabilization is required, the bleed-jet circulation is shut off by a throttle valve. Attached Figure Description

[0024] Figure 1 This is a schematic diagram of an existing compressor using interstage or outlet venting technology for stabilization.

[0025] Figure 2 This is a schematic diagram of the flow distribution framework of the compressor stabilization design method for gas turbines and aero engines provided in the embodiments of this application;

[0026] Figure 3 This is a schematic diagram of the jet channel design in the compressor stabilization design method for gas turbines and aero engines provided in the embodiments of this application;

[0027] Figure 4 This is a schematic diagram illustrating how the compressor stability enhancement design method for gas turbines and aero engines provided in this application shifts the constant speed flow-pressure ratio characteristic to the left, extending the instability boundary.

[0028] Figure 5This is a schematic diagram of the compressor expansion and stabilization design method for gas turbines and aero engines provided in this application, which realizes mass exchange between the circulating jet and the effective inlet airflow through the strong three-dimensional entrainment effect on the blade back side and enables radial transmission of pressure potential energy.

[0029] Figure 6 This is a schematic diagram of the compressor expansion and stability design method for gas turbines and aero engines provided in this application, which suppresses stall backflow on the blade back side, enables the blade basin side to operate under a large incoming flow angle of attack, and improves the blade's power boosting capability.

[0030] Figure 7 This application provides a schematic diagram of a compressor stability enhancement design method for gas turbines and aero engines, which improves the pressure level of the effective outlet airflow of the compressor, increases the compressor near-surge pressure ratio, and expands the compressor instability boundary.

[0031] Figure 8 This is a schematic diagram of a compressor operating in the self-circulating bleed-jet mode provided in the embodiments of this application, where the constant speed characteristic shifts to the left and the matching point changes.

[0032] To better illustrate this embodiment, some parts in the accompanying drawings may be omitted, enlarged, or reduced, and do not represent the actual product dimensions. Furthermore, the drawings are for illustrative purposes only and should not be construed as limiting this application. Detailed Implementation

[0033] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings. Other related parts can be referred to the general design. In the absence of conflict, the embodiments and technical features in the embodiments of this application can be combined with each other to obtain new embodiments.

[0034] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," and "outer," etc., used in this application description to indicate relative direction or positional relationship are used only to indicate relative orientation or positional relationship, and do not imply that the device or component must have a specific orientation, or be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly, and therefore should not be construed as a limitation on this application. The terms "first," "second," "third," and similar terms used in this application description are used only for descriptive purposes to distinguish different components, and should not be construed as indicating or implying relative importance. The terms "a," "one," or "the," etc., used in this application description should not be construed as an absolute limitation on quantity, but should be construed as indicating the existence of at least one. The terms "including," "comprising," etc., used in this application description mean that the element or object preceding the word covers the element or object listed after the word and its equivalents, without excluding other elements or objects.

[0035] Furthermore, it should be noted that, unless otherwise explicitly specified and limited, terms such as “installation,” “connection,” and “linkage” used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium; or it can be a connection within two components. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0036] In multi-stage transonic compressors, under medium and low off-design speed conditions, if the inlet flow coefficient is too low and the blade height range on the inlet stage rotor blade tip side is large, severe stall separation and blade tip leakage will occur, which will couple to form a large-scale channel blockage. This will have varying degrees of negative impact on the flow conditions, matching quality, power boosting level of the compressor front stage, and the stability margin will be low.

[0037] To address the issue of low stability margin of compressors at medium and low off-design speeds, this application provides a compressor stability enhancement design method for gas turbines and aero engines. This method combines the advantages of interstage or outlet bleed technology and micro-flow level self-circulating bleed-jet stability enhancement technology while eliminating their disadvantages. At the same time, it introduces a brand-new stability enhancement mechanism and technology, which can significantly improve the stability margin of multi-stage compressors at medium and low speeds as needed.

[0038] The compressor expansion and stabilization design method for gas turbines and aero engines provided in this application, referencing the flow distribution framework of the self-circulating bleed air-jet expansion and stabilization technology scheme at a small flow rate, designs a bleed air pipeline outside the compressor that allows for large flow rates to pass through, performing bleed air diversion. This diverts a portion of the airflow from the compressor outlet to the inlet position, and reinjects it into the compressor's main flow path in the form of a high-speed jet. Figure 2 As shown, this avoids the large amount of airflow emitted into the environment by the compressor in order to achieve significant expansion and stability, which would take away a large amount of compression energy. This fundamentally ensures that the compressor will not suffer effective power loss related to 100% of the exhaust flow rate, and also avoids the continuous high-frequency noise generated by direct venting.

[0039] The compressor expansion and stabilization design method for gas turbines and aero engines provided in this application, with reference to the amount of venting from interstage or outlet venting technology, rationally designs the bleed air pipeline to ensure that the flow resistance of the bleed air pipeline is small and the cross-sectional area is large enough to allow airflow of up to 10% or more of the compressor inlet flow to return to the compressor inlet position through the bleed air pipeline. Furthermore, a jet channel with circumferential continuous cross-sectional contraction is designed at the compressor inlet position so that the airflow in the bleed air pipeline, driven by the pressure difference between the compressor inlet and outlet, is injected into the main flow channel at high speed in a certain fan-shaped area through the jet channel.

[0040] The compressor stability enhancement design method for gas turbines and aero engines provided in this application involves designing a return flow into the main injection channel, which accounts for a large proportion of the compressor's intake airflow. This significantly reduces the effective flow area of ​​the compressor, causing the compressor's constant speed flow-pressure ratio characteristic to shift to the left, thus expanding the instability boundary. Figure 4 As shown, this expands the local instability boundary.

[0041] The compressor stabilization design method for gas turbines and aero-engines provided in this application involves designing a fan-shaped, high-speed, high-energy return flow into the main jet channel. A small portion of this flow approaches the rotor blade tip casing and flows downstream at high speed, significantly increasing the axial velocity of the incoming flow near the rotor blade tip. This results in a significantly negative blade tip angle of attack, initiating turbine operation. The jet performs work on the rotor blade tip, effectively suppressing strong gap leakage flow. The majority of the jet interacts strongly with the low-energy main jet from upstream in the middle and upper channels of the rotor blades in the inlet stage. On one hand, the strong three-dimensional entrainment effect on the blade back side enables mass exchange between the circulating jet and the effective inlet airflow, and allows for radial transmission of pressure potential energy. Figure 5 As shown, on the other hand, it can strongly suppress stall backflow on the blade back side, allowing the blade basin side to operate at a large incoming flow angle of attack, significantly improving the blade's power boosting capacity, such as... Figure 6 As shown, this increases the pressure level of the effective outlet airflow of the compressor, increases the compressor near-surge pressure ratio, further expands the compressor instability boundary, and fully meets the margin requirements of the compressor during steady / dynamic operation, such as... Figure 7 As shown.

[0042] The compressor operating in the high-flow-rate self-circulating bleed-jet mode involved in this application has its constant speed characteristic shifted to the left, changing the matching point of the entire machine under the same operating conditions. However, based on practical engineering experience, the new matching point will not deviate excessively from the original flow-pressure ratio matching point. Figure 8 As shown in the “new matching zone”, the baseline change for stability margin calculation is relatively limited, and the amplitude of stability expansion achieved by self-circulating bleed air-jet is small.

[0043] The compressor stability design method for gas turbines and aero engines provided in this application involves installing a throttle valve on the bleed air pipeline. This allows for appropriate adjustment of the bleed air-jet circulation flow intensity at different speeds, based on specific usage conditions and stability margin requirements. The basic principle is to minimize the circulation intensity and reduce compressor losses. The position of the "locally extended instability boundary" relative to the original instability boundary is controlled as needed. For the design speed or other higher speeds, the stability margin is usually large enough, and the bleed air-jet circulation flow can be shut off by the throttle valve.

[0044] The compressor expansion and stabilization design method for gas turbines and aero engines provided in this application, in terms of basic flow distribution, is similar to the existing micro- and small-flow-stage self-circulating bleed-jet expansion and stabilization technology schemes, which bleed air from the compressor itself and jet it from itself. In terms of the amount of bleed air into the main flow channel, it is similar to the interstage or outlet venting expansion and stabilization technology schemes, which bleed a large amount of air from itself. However, in terms of the main expansion and stabilization mechanism, it is significantly different from the expansion and stabilization principle of the above-mentioned existing technologies.

[0045] The compressor stability enhancement design method for gas turbines and aero-engines provided in this application involves designing a high-flow-rate bleed-jet system within the multi-stage compressor itself. Bleed air is drawn from the outlet and jetted a certain distance upstream of the inlet stage rotor blade tip for significant stability enhancement at medium and low speeds. The bleed-jet flow rate accounts for a large proportion of the total flow rate. Utilizing the bleed-jet system, the actual flow area within the compressor is reduced, decreasing the actual flow rate at the same speed, significantly expanding the instability boundary. Furthermore, the high-speed, high-energy jet effectively suppresses or even eliminates backflow in the upper channel of the inlet stage rotor, significantly improving rotor operating conditions and enhancing the rotor's power boosting capacity. This increases the compressor outlet airflow pressure ratio and expands the instability boundary, while also improving the overall compressor flow and reducing losses. In addition, the high-energy jet migrates radially from blade tip to blade root within the inlet stage rotor channel, achieving radial migration of high-pressure airflow and radial transport of high-pressure potential energy, further increasing the compressor outlet airflow pressure ratio and expanding the instability boundary. The high-energy jet also performs work on the compressor inlet stage rotor, improving the actual overall operating efficiency.

[0046] The compressor expansion and stabilization design method for gas turbines and aero-engines provided in this application is applied to multi-stage compressors and has significant implications for guiding engineering design. Characterized by a large bleed air / jet flow rate, it can significantly influence the compressor's main flow path and substantially improve the basic low-speed flow framework, thereby achieving substantial expansion and stabilization. This avoids excessive reduction in compressor efficiency caused by direct large-scale bleed air, improves overall thermal cycle efficiency, and eliminates high-frequency noise induced by continuous bleed air, thus improving the human-machine environment. Under actual operating conditions, compared to direct bleed air expansion and stabilization, a large amount of high-energy bleed air impacts the inlet-stage rotor blades in the form of a high-speed jet, performing a certain amount of work on the blades, reducing turbine front-drive shaft work and improving overall efficiency.

[0047] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the scope of protection of this application.

Claims

1. A method for enhancing the stability of a compressor in a gas turbine and aero-engine, characterized in that, include: Set up an air intake pipeline to divert part of the airflow from the compressor outlet to the inlet position and re-inject it into the main flow channel of the compressor in the form of a high-speed jet. The design allows more than 10% of the compressor intake airflow to return to the compressor inlet through the bleed air pipeline; A continuously contracting circumferential jet channel is designed at the compressor inlet. Driven by the pressure difference between the compressor inlet and outlet, the airflow in the intake pipe is injected into the main flow channel at high speed in a fan-shaped area. A small portion of the jet is close to the rotor blade tip casing, where it performs work on the rotor blade tip and suppresses strong gap leakage flow. The majority of the jet interacts with the main flow in the inlet stage rotor blades and above. Through the strong three-dimensional entrainment effect on the blade back side, mass exchange between the circulating jet and the inlet airflow is achieved, and pressure potential energy is radially transferred. At the same time, stall backflow on the blade back side is suppressed, allowing the blade basin side to operate at a large incoming flow angle of attack, thereby improving the blade's power boosting capacity. A throttle valve is installed on the intake air line to meet the requirements for the expansion of the instability boundary and to minimize the cycle intensity and reduce compressor losses, thereby adjusting the intensity of the jet return flow at different speeds.

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